1. Field of the Invention
This invention relates to internally cooled turbine vanes for gas turbine engines and more particularly to the construction of the internally cooled turbine vane comprising a spar and shell construction.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
As one skilled in the gas turbine technology recognizes, the efficiency of the engine is enhanced by operating the turbine at a higher temperature and by increasing the turbine's pressure ratio. Another feature that contributes to the efficiency of the engine is the ability to cool the turbine with a lesser amount of cooling air. The problem that prevents the turbine from being operated at a higher temperature is the limitation of the structural integrity of the turbine component parts that are jeopardized in its high temperature, hostile environment. Scientists and engineers have attempted to combat the structural integrity problem by utilizing internal cooling and selecting high temperature resistant materials. The problem associated with internal cooling is twofold. One, the cooling air that is utilized for the cooling comes from the compressor that has already extended energy to pressurize the air and the spent air in the turbine cooling process in essence is a deficit in engine efficiency. The second problem is that the cooling is through cooling passages and holes that are in the turbine blade or vane which, obviously, adversely affects the blade or vane's structural prowess. Because of the tortuous path (a serpentine path through the blade or vane) that is presented to the cooling air, the pressure drop that is a consequence thereof requires higher supply pressure and more air flow to perform the cooling that would otherwise take a lesser amount of air given the path becomes friendlier to the cooling air. While there are materials that are available and can operate at a higher temperature that is heretofore been used, the problem is how to harness these materials so that they can be used efficaciously in the turbine environment.
To better appreciate these problems it would be worthy of note to recognize that traditional blade cooling approaches include the use of cast nickel based alloys with load-bearing walls that are cooled with radial flow channels and re-supply holes in conjunction with film discharge cooling holes. Examples of these types of blades and vanes are exemplified by the following patents that are incorporated herein by reference.
U.S. Pat. No. 3,378,228 issued to Davies et al on Apr. 16, 1968 shows a blade for a fluid flow duct and comprises ceramic laminations which may be in two or more parts, where the laminations are held together in compression by a hollow tie bar through which cooling air may be passed, and where the blades are mounted between platform members.
U.S. Pat. No. 4,790,721 issued to Morris et al on Dec. 13, 1988 shows an airfoil blade assembly having a metallic core, thin coolant liner and ceramic blade jacket including variable size cooling passages and a circumferential stagnant air gap to provide a substantially cooler core temperature during high temperature operations.
U.S. Pat. No. 4,473,336 issued to Coney et al on Sep. 25, 1984 shows a turbine blade with a spar formed with a central passageway with cooling holes passing through the spar wall into a cavity formed between an airfoil shaped shell and the spar.
U.S. Pat. No. 4,519,745 issued to Rosman et al on May 28, 1985 shows a ceramic blade assembly including a corrugated-metal partition situated in the space between the ceramic blade element and the post member, which corrugated-metal partition forms a compliant layer for the relief of mechanical stresses in the ceramic blade element during aerodynamic and thermal loading of the blade and which partition also serves as a means for defining contiguous sets of juxtaposed passages situated between the ceramic blade element and the post member, one set being open-ended and adjacent to exterior surfaces of the post member for directing cooling fluid there over and the second set being adjacent to the interior surfaces of the ceramic blade element and being closed-off for creating stagnant columns of fluid to thereby insulate the ceramic blade element from the cooling air.
U.S. Pat. No. 4,512,719 issued to Rossmann on Mar. 24, 1981 shows a turbine blade adapted for use with hot gases comprising a radially inward portion of metal including a core projecting radially outwards on which is supported a ceramic portion of airfoil section enclosing the core. The inner end of the ceramic portion forms a continuous surface contour with the metal inward portion. The ceramic portion extends no more than one-half of the total span of the blade and, preferably, about one-third of the blade span. In a particular embodiment, the wall thickness of the ceramic portion can increase in a radially outwards direction.
U.S. Pat. No. 4,563,128 issued to Rossmann on Jan. 7, 1986 shows a hot gas impinged turbine blade suitable for use under super-heated gas operating conditions has a hollow ceramic blade member and an inner metal support core extending substantially radially through the hollow blade member and having a radially outer widened support head. The support head has radially inner surfaces against which the ceramic blade member supports itself in a radial direction on both sides of the head. The radially inner surfaces of the head are inclined at an angle to the turbine axis so as to form a wedge or key forming a dovetail type connection with respectively inclined surfaces of the ceramic blade member. This dovetail type connection causes a compressive stress on the ceramic blade member during operation, whereby an optimal stress distribution is achieved in the ceramic blade member.
U.S. Pat. No. 4,247,259 issued to Saboe et al on Jan. 27, 1981 shows a composite, ceramic/metallic fabricated blade unit for an axial flow rotor includes an elongated metallic support member having an airfoil-shaped strut, one end of which is connected to a dovetail root for attachment to the rotor disc, while the opposite end thereof includes an end cap of generally airfoil-shape. The circumferential undercut extending between the end cap and the blade root is clad with an airfoil-shaped ceramic member such that the cross-section of the ceramic member substantially corresponds to the airfoil-shaped cross-section of the end cap, whereby the resulting composite ceramic/metallic blade has a smooth, exterior airfoil surface. The metallic support member has a longitudinally extending opening through which coolant is passed during the fabrication of the blade. Simultaneously, ceramic material is applied and bonded to the outer surface of the elongated airfoil-shaped strut portion, with the internal cooling of the metallic strut during the processing operation allowing the metal to withstand the processing temperature of the ceramic material.
U.S. Pat. No. 3,694,104 issued to Erwin on Sep. 26, 1972 shows a turbomachinery blade secured to a rotor disc by a pin.
U.S. Pat. No. 4,314,794 issued to Holden, deceased et al on Feb. 9, 1982 shows a transpiration cooled blade for a gas turbine engine is assembled from a plurality of individual airfoil-shaped hollow ceramic washers stacked upon a ceramic platform which in turn is seated on a metal root portion. The airfoil portion so formed is enclosed by a metal cap covering the outermost washer. A metal tie tube is welded to the cap and extends radially inwardly through the hollow airfoil portion and through aligned apertures in the platform and root portion to terminate in a threaded end disposed in a cavity within the root portion housing a tension nut for engagement thereby. The tie tube is hollow and provides flow communication for a coolant fluid directed through the root portion and into the hollow airfoil through apertures in the tube. The ceramic washers are made porous to the coolant fluid to cool the blade via transpiration cooling.
U.S. Pat. No. 3,644,060 issued to Bryan on Feb. 22, 1972 shows a cooled airfoil in which a shell is secured over a spar by dove-tail grooves.
U.S. Pat. No. 4,257,737 issued to Andress et al on Apr. 23, 1985 shows a Cooled Rotor Blade, where the cooled rotor blade is constructed having a cooling passage extending from the root and through the airfoil shaped section in a serpentine fashion, making several passes between the bottom and top thereof; a plurality of openings connect said cooling passage to the trailing edge; a plurality of compartments are formed lengthwise behind the leading edge of the blade; said compartments having openings extending through to the exterior forward portion of the blade; and sized openings connect the cooling passage to each of the compartments to control the pressure in each compartment.
U.S. Pat. No. 4,753,575 issued to Levengood et al on Jun. 28, 1988 shows an airfoil with nested cooling channels, where the hollow, cooled airfoil has a pair of nested, coolant channels therein which carry separate coolant flows back and forth across the span of the airfoil in adjacent parallel paths. The coolant in both channels flows from a rearward to forward location within the airfoil allowing the coolant to be ejected from the airfoil near the leading edge through film coolant holes.
U.S. Pat. No. 5,476,364 issued to Kildea on Dec. 19, 1995 shows a tip seal and anti-contamination for turbine blades, where a cavity is judiciously dimensioned and located adjacent the tip's surface discharge port of internally cooling passage of the airfoil of the turbine blade of a gas turbine engine and extending from the pressure surface to the back wall of the discharge port guards against the contamination and plugging of the discharge port.
U.S. Pat. No. 5,700,131 issued to Hall et al on Dec. 23, 1997 shows an internally cooled turbine blade for a gas turbine engine that is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length. Trip strips may be employed to augment the pressure drop distribution.
Also well known by those skilled in this technology is that the engine's efficiency increases as the pressure ratio of the turbine increases and the weight of the turbine decreases. Needless to say, these parameters have limitations. Increasing the speed of the turbine also increases the airfoil loading and, of course, satisfactory operation of the turbine is to stay within given airfoil loadings. The airfoil loadings are governed by the cross sectional area of the turbine multiplied by the velocity of the tip of the turbine squared, or AN2. Obviously, the rotational speed of the turbine has a significant impact on the loadings.
The spar/shell construction contemplated by this invention affords the turbine engine designer the option of reducing the amount of cooling air that is required in any given engine design. And in addition, allowing the designer to fabricate the shell from exotic high temperature materials that heretofore could not be cast or forged to define the surface profile of the airfoil section. In other words, by virtue of this invention, the shell can be made from Niobium or Molybdenum or their alloys, where the shape is formed by a well known electric discharge process (EDM) or wire EDM process. In addition, because of the efficacious cooling scheme of this invention, the shell-portion could be made from ceramics, or more conventional materials and still present an advantage to the designer because a lesser amount of cooling air would be required.